U.S. patent application number 15/534393 was filed with the patent office on 2018-09-13 for improved directional control for coaxial rotary wing craft.
This patent application is currently assigned to Sikorsky Aircraft Corporation. The applicant listed for this patent is Sikorsky Aircraft Corporation. Invention is credited to Mark W. Scott.
Application Number | 20180257771 15/534393 |
Document ID | / |
Family ID | 56564477 |
Filed Date | 2018-09-13 |
United States Patent
Application |
20180257771 |
Kind Code |
A1 |
Scott; Mark W. |
September 13, 2018 |
IMPROVED DIRECTIONAL CONTROL FOR COAXIAL ROTARY WING CRAFT
Abstract
A rotary winged aircraft includes an airframe and a drive system
located at the airframe. A main rotor system is positioned at the
airframe and is operably connected to the drive system to provide
lift for the rotary winged aircraft. An auxiliary propulsor is
located at the airframe and includes a plurality of propeller
blades rotatable about a propulsor axis. Collective and cyclic
pitch input applied to the auxiliary propeller blades increases yaw
performance of the aircraft. A method of operating a rotary wing
aircraft includes powering an auxiliary propulsor secured to an
airframe of the aircraft and including a plurality of propeller
blades rotatable about a propulsor axis. Individual propeller
blades are cyclically rotated about their respective propeller
blade axes to cyclically change a propeller blade pitch. Rotation
of the aircraft about a yaw axis is induced via the collective and
cyclic pitch change of the propeller blades.
Inventors: |
Scott; Mark W.; (Bethany,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Sikorsky Aircraft Corporation |
Stratford |
CT |
US |
|
|
Assignee: |
Sikorsky Aircraft
Corporation
Stratford
CT
|
Family ID: |
56564477 |
Appl. No.: |
15/534393 |
Filed: |
November 23, 2015 |
PCT Filed: |
November 23, 2015 |
PCT NO: |
PCT/US2015/062158 |
371 Date: |
June 8, 2017 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
62111748 |
Feb 4, 2015 |
|
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
B64C 2027/8281 20130101;
B64C 2027/8236 20130101; B64C 27/10 20130101; B64C 2027/8272
20130101; B64C 27/82 20130101 |
International
Class: |
B64C 27/82 20060101
B64C027/82 |
Goverment Interests
FEDERAL RESEARCH STATEMENT
[0001] This invention was made with government support with the
United States Army under Contract No. W911W6-13-2-0003. The
government therefore has certain rights in this invention.
Claims
1. A rotary winged aircraft comprising: an airframe; a drive system
disposed at the airframe; a main rotor system disposed at the
airframe and operably connected to the drive system to provide lift
for the rotary winged aircraft; and an auxiliary propulsor disposed
at the airframe including a plurality of propeller blades rotatable
about a propulsor axis, such that a combination of collective and
cyclic pitch input applied to the auxiliary propeller blades
increases yaw performance of the aircraft.
2. The aircraft of claim 1, wherein the plurality of propeller
blades are operably connected to a propeller hub.
3. The aircraft of claim 2, wherein the cyclic pitch input results
in a moment about the propeller hub.
4. The aircraft of claim 1, wherein the collective and cyclic pitch
input results in a net force perpendicular to the propulsor
axis.
5. The aircraft of claim 1, wherein the collective and cyclic pitch
input are applied during one of hover flight, low speed flight
conditions, low rotor thrust conditions or during autorotational
landing.
6. The aircraft of claim 1, wherein the main rotor system is dual
coaxial rotor system.
7. The aircraft of claim 1, wherein the auxiliary propulsor is
disposed at an extended tail of the airframe.
8. A method of operating a rotary wing aircraft comprising:
powering an auxiliary propulsor secured to an airframe of the
aircraft and including a plurality of propeller blades rotatable
about a propulsor axis; rotating the plurality of propeller blades
about the propulsor axis; cyclically rotating individual propeller
blades about their respective propeller blade axes to cyclically
change a propeller blade pitch; and inducing rotation of the
aircraft about a yaw axis via the collective and cyclic pitch
change of the propeller blades.
9. The method of claim 8, wherein the plurality of propeller blades
are operably connected to a propeller hub.
10. The method of claim 9, wherein the cyclic pitch change results
in a moment about the propeller hub.
11. The method of claim 8, wherein the collective and cyclic pitch
change results in a net force perpendicular to the propulsor
axis.
12. The method of claim 8, further comprising applying the
collective and cyclic pitch change during one of hover flight, low
speed flight conditions, low rotor thrust conditions or during
autorotational landing.
13. An auxiliary propulsor for a rotary winged aircraft comprising:
a propeller hub; and a plurality of propeller blades rotatable
about a propulsor axis, such that a cyclic pitch change applied to
the auxiliary propeller blades increases yaw performance of the
aircraft.
14. The auxiliary propulsor of claim 13, wherein the cyclic pitch
change results in a moment about the propeller hub.
15. The auxiliary propulsor of claim 13, wherein collective and
cyclic pitch change results in a net force perpendicular to the
propulsor axis.
16. The auxiliary propulsor of claim 13, wherein the collective and
cyclic pitch change is applied during one of hover flight, low
speed flight conditions, low rotor thrust conditions or during
autorotational landing.
17. The auxiliary propulsor of claim 13, wherein the auxiliary
propulsor is disposed at an extending tail of the rotary wing
aircraft.
Description
BACKGROUND
[0002] The subject matter disclosed herein generally relates to
directional control of aircraft. More specifically, the subject
disclosure relates to yaw control of dual coaxial rotor rotary wing
aircraft.
[0003] A rotary wing aircraft, or helicopter, with a coaxial
contra-rotating rotor system is capable of higher speeds as
compared to conventional single rotor helicopters due in part to
the balance of lift between advancing sides of the main rotor
blades on the upper and lower rotor systems. Such aircraft,
however, tend to have weak yaw control during hover flight, low
speed flight conditions, under low rotor thrust conditions and
during autorotational landings. Low speed/low rotor thrust flight
conditions are encountered in aircraft return to target and similar
"low g" maneuvers. The ability to safely autorotate is an important
feature, especially for single engine aircraft. The poor yaw
performance is typically due to the inability of the aircraft to
create adequate differential rotor torque between the two coaxial
rotors under these flight conditions.
BRIEF SUMMARY
[0004] In one embodiment, a rotary winged aircraft includes an
airframe and a drive system located at the airframe. A main rotor
system is positioned at the airframe and is operably connected to
the drive system to provide lift for the rotary winged aircraft. An
auxiliary propulsor is located at the airframe and includes a
plurality of propeller blades rotatable about a propulsor axis. A
cyclic pitch input applied to the auxiliary propeller blades
increases yaw performance of the aircraft.
[0005] Additionally or alternatively, in this or other embodiments
the plurality of propeller blades are operably connected to a
propeller hub.
[0006] Additionally or alternatively, in this or other embodiments
the cyclic pitch input results in a moment about the propeller
hub.
[0007] Additionally or alternatively, in this or other embodiments
cyclic and collective pitch inputs result in a net force
perpendicular to the propulsor axis.
[0008] Additionally or alternatively, in this or other embodiments
the cyclic and collective pitch inputs are applied during hover
flight, low speed flight conditions, low rotor thrust conditions,
or during autorotational landing.
[0009] Additionally or alternatively, in this or other embodiments
the main rotor system is dual coaxial rotor system.
[0010] Additionally or alternatively, in this or other embodiments
the auxiliary propulsor is disposed at an extending tail of the
airframe.
[0011] A method of operating a rotary wing aircraft includes
powering an auxiliary propulsor secured to an airframe of the
aircraft and including a plurality of propeller blades rotatable
about a propulsor axis. The plurality of propeller blades are
rotated about the propulsor axis. Individual propeller blades are
cyclically rotated about their respective propeller blade axes to
cyclically change a propeller blade pitch. Rotation of the aircraft
about a yaw axis is induced via the cyclic pitch change of the
propeller blades.
[0012] Additionally or alternatively, in this or other embodiments
the plurality of propeller blades are operably connected to a
propeller hub.
[0013] Additionally or alternatively, in this or other embodiments
the cyclic pitch change results in a moment about the propeller
hub.
[0014] Additionally or alternatively, in this or other embodiments
the cyclic and collective pitch change results in a net force
perpendicular to the propulsor axis.
[0015] Additionally or alternatively, in this or other embodiments
the cyclic pitch change is applied during hover flight, low speed
flight conditions, low rotor thrust conditions, or during
autorotational landing.
[0016] An auxiliary propulsor for a rotary winged aircraft includes
a propeller hub and a plurality of propeller blades rotatable about
a propulsor axis, such that a cyclic pitch change applied to the
auxiliary propeller blades increases yaw performance of the
aircraft.
[0017] Additionally or alternatively, in this or other embodiments
the cyclic pitch change results in a moment about the propeller
hub.
[0018] Additionally or alternatively, in this or other embodiments
the cyclic and collective pitch change results in a net force
perpendicular to the propulsor axis.
[0019] Additionally or alternatively, in this or other embodiments
the cyclic pitch change is applied during hover flight, low speed
flight conditions, low rotor thrust conditions or during
autorotational landing.
[0020] Additionally or alternatively, in this or other embodiments
the auxiliary propulsor is located at an extended tail of the
rotary wing aircraft.
[0021] These and other advantages and features will become more
apparent from the following description taken in conjunction with
the drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
[0022] The subject matter, which is regarded as the invention, is
particularly pointed out and distinctly claimed in the claims at
the conclusion of the specification. The foregoing and other
features, and advantages of the invention are apparent from the
following detailed description taken in conjunction with the
accompanying drawings in which:
[0023] FIG. 1 is a general side view of an exemplary rotary wing
aircraft for use in accordance with embodiments;
[0024] FIG. 2 is a plan view of an embodiment of a rotary winged
aircraft; and
[0025] FIG. 3 is a plan view of another embodiment of a rotary
winged aircraft.
[0026] The detailed description explains embodiments of the
invention, together with advantages and features, by way of example
with reference to the drawings.
DETAILED DESCRIPTION
[0027] FIG. 1 illustrates an exemplary vertical takeoff and landing
(VTOL) high speed compound or coaxial contra-rotating rigid rotor
aircraft 10 having a dual, contra-rotating main rotor system 12,
which rotates about a rotor axis 36. The aircraft 10 includes an
airframe 14 which supports the dual, contra-rotating, coaxial main
rotor system 12 as well as a translational thrust system 16 which
provides translational thrust generally parallel to an aircraft
longitudinal axis 18.
[0028] The main rotor system 12 includes an upper rotor system 20
and a lower rotor system 22 as dual counter rotating main rotors in
a coaxial configuration. A plurality of rotor blade assemblies 24
is mounted to a rotor hub assembly 26 of each rotor system 20, 22.
The main rotor system 12 is driven by a transmission 28. The
translational thrust system 16 may be any propeller system
including, but not limited to a pusher propeller, a tractor
propeller, a nacelle mounted propeller, etc. In the example of FIG.
1, the translational thrust system 16 includes an auxiliary
propulsor 30. In an embodiment, the auxiliary propulsor 30 is a
pusher propeller system with a propeller rotational axis oriented
substantially horizontal and parallel to the aircraft longitudinal
axis 18 to provide thrust for high speed flight. The translational
thrust system 16 may be driven through a main gearbox 32 which also
drives the main rotor system 12.
[0029] The transmission 28 includes the main gearbox 32 driven by
one or more engines, illustrated schematically at 34. The main
gearbox 32 and engines 34 are considered as part of the
non-rotating frame of the aircraft 10. In the case of a rotary wing
aircraft, the main gearbox 32 may be interposed between one or more
gas turbine engines 34, the main rotor system 12 and the
translational thrust system 16. In one embodiment, the main gearbox
32 is a split torque gearbox which carries torque from the engines
34 through a multitude of drivetrain paths. Although a particular
rotary wing aircraft configuration is illustrated and described in
the disclosed non-limiting embodiment, other configurations and/or
machines with rotor systems are within the scope of the present
invention. It is to be appreciated that while the description
herein relates to a rotary wing aircraft with a dual coaxial
contra-rotating rotor system, the disclosure herein may be as
readily applied to other rotor systems, such as turboprops,
tilt-rotors, and tilt-wing aircraft, or a conventional single rotor
system.
[0030] As stated above, typical aircraft with coaxial counter
rotating main rotors have weak yaw control in certain operating
conditions. Referring now to FIGS. 2 and 3, in embodiments of the
present disclosure, cyclic pitch control is applied to propeller
blades 40 of the auxiliary propulsor 30 to improve yaw response
about a yaw axis 42 of the aircraft 10 in selected flight
conditions including, but not limited to, hover flight, low speed
flight conditions, under low rotor thrust conditions, and during
autorotational landings. Low speed/low rotor thrust flight
conditions are encountered in aircraft return to target and similar
"low g" maneuvers. In applying cyclic pitch control, a pitch about
a propeller blade axis 44 of each propeller blade 40 is varied as
the propeller blade 40 travels about the propeller rotational axis
46.
[0031] Referring now to FIG. 2, a symmetric cyclic pitch change
input of approximately 10 degrees is utilized to create a pure
moment around a propeller hub 48 of the auxiliary propulsor 30 as
shown. This moment results in an increase to overall aircraft 10
yaw moment about the yaw axis 42. Thus increasing the yaw rate and
overall control.
[0032] Referring now to FIG. 3, cyclic and collective pitch input
are applied to blades 40 that create flat pitch (near 0 degrees) of
propeller blades 40 at one side of the auxiliary propulsor 30, and
a coarse pitch of 70 degrees or more at an opposing side of the
auxiliary propulsor 30. Collective pitch application in combination
with cyclic pitch application is required to create coarse pitch on
one half of the propulsor and flat pitch on the other. The coarse
pitch blades on one half of the propulsor 30 creates very high drag
because they are stalled which far exceeds the blade drag force on
the opposite side of the propulsor 30 at flat pitch. The result is
a substantial net force perpendicular to the propeller rotational
axis 46 as shown in FIG. 3. The magnitude of force depends on the
difference in pitch between propeller blades 40 on the opposing
sides of the auxiliary propulsor 30, while the direction of the
force depends on which side of the auxiliary propulsor 30 the
coarse pitch is applied to. The large force perpendicular to the
rotational axis 46 at a distance from the yaw axis 42 creates the
additional aircraft yawing moment.
[0033] The difference between the modes of FIGS. 2 and 3 is a
magnitude of the pitch change. Determination of the best mode
depends on particular aircraft and propulsor characteristics such
as propulsor distance from the rotor axis 42, propulsor diameter,
blade area, and tip speed. Aircraft design yaw control requirements
also has impact on selection on the best mode of operation.
[0034] Utilizing cyclic pitch input at the auxiliary propulsor 40
results in increased yaw control of the aircraft 10 throughout its
flight envelope. This disclosure eliminates the need for ancillary
devices, such as an auxiliary tail rotor, which are heavy and
increase drag on the aircraft. It also eliminates the need for
rotor blown control surfaces that are not only heavy and draggy but
also of limited value in low speed and hovering flight due to low
velocity airflow.
[0035] While the invention has been described in detail in
connection with only a limited number of embodiments, it should be
readily understood that the invention is not limited to such
disclosed embodiments. For instance, while described in terms of
aircraft, it is understood that aspects could be used in other
contexts such as for wind turbines, maritime propulsion, or other
technologies in which a rotating element's plane of rotation will
vary from perpendicular with the axis of rotation of a shaft
driving the rotating element. Rather, the invention can be modified
to incorporate any number of variations, alterations, substitutions
or equivalent arrangements not heretofore described, but which are
commensurate with the spirit and scope of the invention.
Additionally, while various embodiments of the invention have been
described, it is to be understood that aspects of the invention may
include only some of the described embodiments. Accordingly, the
invention is not to be seen as limited by the foregoing
description, but is only limited by the scope of the appended
claims.
* * * * *